CN113803164A - Ultra-miniature turbojet power device - Google Patents
Ultra-miniature turbojet power device Download PDFInfo
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- CN113803164A CN113803164A CN202111087116.XA CN202111087116A CN113803164A CN 113803164 A CN113803164 A CN 113803164A CN 202111087116 A CN202111087116 A CN 202111087116A CN 113803164 A CN113803164 A CN 113803164A
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- turbine
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- combustion chamber
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
- F02C3/08—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising at least one radial stage
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/14—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
- F02C3/16—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant the combustion chambers being formed at least partly in the turbine rotor or in an other rotating part of the plant
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/22—Fuel supply systems
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The invention discloses an ultra-miniature turbojet power device. The device comprises: the compressor rotor of the single-stage radial-flow compressor and the turbine rotor of the single-stage radial turbine form an integral rotor; a microscale combustion chamber is formed between the single-stage radial-flow compressor and the single-stage radial-center turbine, and the microscale combustion chamber comprises: the high-temperature and high-pressure gas flows into a driving turbine rotor of the turbine along the single-stage radial direction to do work, and then is converted into axial discharge again; the turbojet power device is formed by adopting 6 monocrystalline silicon wafers with the thickness of 0.4mm, 2 monocrystalline silicon wafers with the thickness of 0.6mm and 3 monocrystalline silicon wafers with the thickness of 0.8mm to be etched and formed separately and then sealing the adjacent silicon wafers into an air flow channel through a diffusion bonding connection technology. The invention can meet the urgent requirements of the micro unmanned aerial vehicle on the power-weight ratio of the power device, and has larger potential for improving the performance.
Description
Technical Field
The invention relates to the technical field of micro turbine engines, in particular to an ultra-micro turbine jet power device.
Background
The Micro unmanned aerial Vehicle (MAV) has the characteristics of small volume, high speed, strong detection capability, good concealment, simple operation, flexibility and the like, is used as a Micro reconnaissance aircraft in the military field for battlefield reconnaissance, target monitoring, target searching, signal interference, close range combat, even reconnaissance of important sensitive parts of enemies in large-scale building facilities and interception of enemy situations, and can also be used for detecting military dynamic tasks in dangerous areas which are not easy to reach by human beings or are not suitable for being involved, such as biochemical pollution areas, nuclear radiation dangerous areas and the like, thereby adding 'eyes in the Air' to the battlefield; in the civil aspect, the MAV is mainly used for environmental monitoring, meteorological monitoring, forest fire prevention monitoring and flood disaster monitoring, and can also be used in the fields of aerial photography, crop monitoring, pesticide spraying and the like. At present, western military strong countries enter a research hot tide stage of micro unmanned aerial vehicles, but the most important and most urgent core technology to be solved for developing the micro unmanned aerial vehicles is the problem of the ultramicro power device with high energy density, so that the research on novel ultramicro power devices with high energy storage density, small size, light weight and high power is urgent.
An Ultra-Micro Turbine jet power plant (UMTE) has the characteristics of high power density and high energy, and is a high-energy-density power plant urgently needed by a Micro unmanned aerial vehicle. According to the special thermal cycle and structural characteristics of UMTE, the power device needs to adopt a structural form that a single-stage radial-flow compressor, a microscale combustion chamber and a single-stage radial-center turbine are integrated, the number of parts and the number of connecting points are as small as possible, and the diameter of the whole device is no more than 22mm, and the axial length is no more than 6 mm; because of the restriction of small flow, small size, high rotating speed and high power density, the pressure ratio of the gas flow in the single-stage compressor is not less than 4, the tangential speed of the rotor reaches 500m/s, the residence time of the gas in the combustion chamber is only 0.5ms, which is one order of magnitude lower than that of the gas flow in the traditional combustion chamber, and the temperature of the gas flowing into the turbine is not less than 1500K.
Disclosure of Invention
The technical problem solved by the invention is as follows: overcomes the defects of the prior art and provides an ultra-miniature turbojet power device.
In order to solve the above-mentioned problems, an embodiment of the present invention provides an ultra-miniature turbojet power device, including:
the device comprises: a single-stage radial-flow compressor, a microscale combustion chamber and a single-stage radial-inflow turbine, wherein,
the compressor rotor of the single-stage radial-flow compressor and the turbine rotor of the single-stage radial-flow turbine form an integral rotor;
the single-stage radial-flow compressor and the single-stage radial-inflow turbine form the micro-scale combustion chamber therebetween, and the micro-scale combustion chamber comprises: the high-temperature and high-pressure gas flows into a driving turbine rotor of the turbine along the single-stage radial direction to do work, and then is converted into axial discharge again;
the turbojet power device is formed by adopting 6 monocrystalline silicon wafers with the thickness of 0.4mm, 2 monocrystalline silicon wafers with the thickness of 0.6mm and 3 monocrystalline silicon wafers with the thickness of 0.8mm to be etched and formed separately and then sealing the adjacent silicon wafers into an air flow channel through a diffusion bonding connection technology.
Optionally, the single-stage radial compressor is composed of an air inlet channel, a compressor rotor, a compressor stator and a fuel injection port; wherein the content of the first and second substances,
the air inlet channel is arranged at the center of the compressor stator;
the compressor rotor is provided with a plurality of blades, so that the airflow passing through the air inlet channel flows in the radial direction after passing through the compressor rotor and then passes through diffuser blades on the compressor stator to play a role in speed reduction and diffusion;
the fuel injection port is arranged on the section of the stator outlet of the compressor in an annular structure, and is in a uniform multipoint form in the circumferential direction.
Optionally, the micro-scale combustion chamber is composed of an annular cavity heat insulation interlayer, a flame tube, an ignition electric nozzle, a protruding flaring and a fuel gas outlet, wherein,
the annular cavity heat insulation interlayer is arranged on the outer side of the micro-scale combustion chamber in an annular cylindrical shape so as to cool the chamber wall of the combustion chamber and preheat premixed gas;
the protruding flaring is arranged at the position of the airflow inlet of the flame tube, and the ignition electric nozzles are circumferentially arranged in the front of the combustion chamber so as to ignite the combustion chamber to organize and burn, and then flow into the single-stage radial inflow turbine through the gas outlet.
Optionally, the single stage centripetal turbine is comprised of a turbine guide, a turbine rotor and an exhaust outlet, wherein,
the turbine guider is provided with a plurality of guide blades, and after high-temperature airflow passes through the turbine guider, the high-temperature airflow flows into the blades of the turbine rotor along the radial direction to drive the turbine rotor to do work, and then the high-temperature airflow is bent by 90 degrees and is axially discharged from the annular exhaust passage.
Optionally, the compressor stator has an outlet diameter of 7mm, and the fuel injection port is circumferentially disposed at the outlet.
Optionally, the width of the annular cavity heat insulation interlayer is 0.4mm, and the axial length is 4.4 mm.
Compared with the prior art, the invention has the advantages that:
the design inlet flow of the ultra-micro turbojet power device is 0.35g/s, the thrust is 0.3N, the temperature in front of the turbine is 1500K, the urgent requirements of the micro unmanned aerial vehicle on the power-weight ratio of the power device can be met, and meanwhile, the potential for improving the performance is large; and the ultra-micro turbine jet power device has compact structure, small size (diameter is 22mm, length is 6mm) and light weight (4.2g), and can meet the structural requirements of micro unmanned aerial vehicle power.
Drawings
Fig. 1 is a schematic structural diagram of an ultra-miniature turbojet power unit according to an embodiment of the present invention;
fig. 2 is a schematic structural diagram of a single-stage radial-flow compressor provided in an embodiment of the present invention;
FIG. 3 is a schematic structural diagram of a micro-scale combustion chamber according to an embodiment of the present invention;
FIG. 4 is a schematic structural diagram of a single-stage radial inflow turbine provided in accordance with an embodiment of the present invention;
fig. 5 is a schematic structural diagram of an integral rotor according to an embodiment of the present invention.
Detailed Description
The scheme adopts the small-flow radial-flow high-pressure-ratio compressor, simultaneously arranges annular fuel injection points along the outlet flow channel of the compressor in advance, and arranges an annular cavity heat insulation interlayer outside the combustion chamber, so that the injected fuel is premixed with main flow as early as possible, and premixed gas can be preheated, thereby being beneficial to prolonging the mixing time, forming uniformly-distributed gas, and designing a protruding flaring at the inlet of the combustion chamber to improve the flame stability.
Next, the technical solutions of the embodiments of the present invention will be described in detail below with reference to the drawings attached to the specification.
Referring to fig. 1, there is shown a schematic structural view of an ultra-miniature turbojet power unit according to an embodiment of the present invention, and as shown in fig. 1, the ultra-miniature turbojet power unit may include: the radial-flow type single-stage radial-flow type gas compressor comprises a single-stage radial-flow type gas compressor 1, a micro-scale combustion chamber 2, a single-stage radial-flow type turbine 3 and an integral rotor 4.
As shown in fig. 1 to 5, the compressor rotor 1b of the radial compressor 1 and the turbine rotor 3b of the radial turbine 3 are designed as an integral rotor 4; a micro-scale combustion chamber 2 is designed between the radial-flow type compressor 1 and the centripetal turbine 3, the micro-scale combustion chamber 2 is designed into a flame tube 2b with an annular fuel jet orifice 1d, an annular cavity heat insulation interlayer 2a and a sudden expansion port 2d to ensure that the fuel and the main flow are premixed and preheated fully, simultaneously, the tempering can be prevented, the flame is stabilized, the retention time of the gas in the combustion chamber is prolonged, and finally, the high-temperature and high-pressure gas flows into the turbine 3 along the radial direction to drive a turbine rotor 3b to do work and then is converted into axial discharge again;
the single-stage radial-flow compressor 1 consists of an air inlet channel 1a, a compressor rotor 1b, a compressor stator 1c and a fuel injection port 1 d; the air inlet 1a is arranged at the center of the compressor stator 1c, 20 blades 1b-1 with the blade height of 0.58mm are designed on the compressor rotor 1b, so that air flow passing through the air inlet 1a flows through the compressor rotor 1b along the radial direction and then passes through 13 diffuser blades 1c-1 with the blade height of 0.6mm designed on the compressor stator 1c for speed reduction and pressure expansion, and meanwhile, in order to ensure that air is fully mixed in a limited space, 15 fuel injection ports 1d with the diameter of 0.3mm are uniformly arranged on the outlet R of the compressor stator 1b in the circumferential direction of a section with the diameter of 7 mm;
the micro-scale combustion chamber 2 consists of a combustion chamber heat insulation interlayer 2a, a flame tube 2b, an ignition electric nozzle 2c, a sudden expansion port 2d and a fuel gas outlet 2 e; the combustor interlayer 2a is 0.4mm wide and 4.4mm long in axial length, the annular cavity is arranged on the outer side of the combustor 2 to play a role in cooling the wall of the combustor and preheating premixed gas, meanwhile, airflow is uniformly distributed in the annular cavity heat-insulating interlayer after being premixed for 1.4ms, a 0.6mm wide sudden expansion port 2d is designed at the inlet of the combustor to achieve the purposes of stabilizing flame and prolonging the residence time of the gas, and 6 ignition electric nozzles 2c are uniformly designed in the flame tube 2b in the circumferential direction of the 9mm section;
the single-stage radial inflow turbine 3 is composed of a turbine guide 3a, a turbine rotor 3b, and an exhaust port 3 c; 24 guide blades 3a-1 with the blade height of 0.3mm are designed on the turbine guide device 3a, and high-temperature airflow flows into 15 turbine rotor blades 3b-1 with the blade height of 0.27mm after passing through the guide device 3a, drives the turbine rotor 3b to do work, then turns 90 degrees, and is axially discharged from the annular exhaust passage 3 c.
The implementation process of the ultra-micro turbojet power device provided by the embodiment of the invention is given below, and specifically comprises the following steps:
step 1: according to the special thermal cycle and structural characteristics of UMTE, namely flow 0.35g/s, fuel flow 0.01g/s, diameter no more than 22mm, axial length 6mm, rotor tangential speed up to 500m/s, turbine inlet gas temperature not lower than 1500K, a single-stage radial flow type air compressor, a micro-scale combustion chamber and a single-stage centripetal turbine are integrated into a whole, the whole machine is limited by structural size, a micro-electro-mechanical system (MEMS) manufacturing technology is adopted, a cavity, a fixed blade, a rotating blade, a fuel nozzle and an air inlet/outlet of the device are independently etched on a single-layer silicon wafer by using allowable temperature up to 1500 ℃, 6 single-layer silicon wafers with thickness of 0.4mm, 2 single-layer silicon wafers with thickness of 0.6mm and 3 single-layer centripetal turbines with thickness of 0.8mm, a semiconductor micro-etching technology is adopted, and the cavity, the fixed blade, the rotating blade, the fuel nozzle and the air inlet/outlet are connected with the single-layer silicon layer to form a whole by using a diffusion bonding technology.
Step 2: under the attraction of 20 compressor rotor blades with the blade height of 0.58mm, airflow flows into the air inlet, turns 90 degrees and enters a flow channel with the height of 0.6mm, the pressure ratio of the pressurized air compressor reaches 4.0, and then the pressurized air compressor is subjected to deceleration diffusion through the blades (13 blades with the blade height of 0.6mm) of the diffuser of the air compressor and then turns 90 degrees and flows into a combustion chamber;
and step 3: in order to ensure that the gas is completely and fully mixed in a limited space, 15 fuel injection ports with the diameter of 0.3mm are uniformly designed in the circumferential direction of the section of the outlet R (7.0 mm) of the stator of the gas compressor, so that the fuel is mixed with the main flow in advance;
and 4, step 4: designing an annular cavity heat insulation interlayer with the width of 0.4mm and the length of 4.4mm at the position of 0.3mm between the outermost side of the combustion chamber and the flame tube, further uniformly mixing air flow in the annular cavity heat insulation interlayer, simultaneously playing a cooling protection role on the flame tube of the combustion chamber, turning the air flow by 180 degrees to enter a 0.6mm protruded flaring of the combustion chamber, enabling the air flow velocity passing through the protruded flaring to be about 5.0m/s, enabling the mixed gas to flow into the combustion chamber, then uniformly designing 6 ignition electric nozzles in the circumferential direction through the section R which is 9mm, enabling ignition tissues to burn, and enabling the mixed gas to stay for about 3.5ms and then enter a turbine;
and 5: after being rectified and accelerated by 24 guide vanes with the blade height of 0.3mm, the fuel gas with the temperature of about 1500K flows into 15 turbine rotor blades with the blade height of 0.27mm along the radial direction to drive the whole rotor to work, and finally is discharged from an exhaust port in the axial direction after rotating 90 degrees to generate reverse thrust.
The ultra-miniature turbine jet power device provided by the embodiment of the invention is based on micro-electro-mechanical etching technology of a monocrystalline silicon wafer, effectively organizes small-flow gas distribution by breaking through an integrated design method of a radial-flow type gas compressor, a microscale combustion chamber and a centripetal turbine, reasonably arranges a geometric space of a small-size structure, and meets the requirements of high-power, small-size and light-weight miniature unmanned aerial vehicle power. Its main advantages and effects are as follows:
1) the designed inlet flow of the ultra-micro turbojet power device is 0.35g/s, the thrust is 0.3N, the temperature in front of the turbine is 1500K, the urgent requirements of the micro unmanned aerial vehicle on the power-weight ratio of the power device can be met, and meanwhile, the potential for improving the performance is large;
2) the ultra-micro turbojet power device is compact in structure, small in size (the diameter is 22mm, the length is 6mm), light in weight (4.2g), and capable of meeting the structural requirements of power of a micro unmanned aerial vehicle.
The detailed description set forth herein may provide those skilled in the art with a more complete understanding of the present application, and is not intended to limit the present application in any way. Thus, it will be appreciated by those skilled in the art that modifications or equivalents may still be made to the present application; all technical solutions and modifications thereof which do not depart from the spirit and technical essence of the present application should be covered by the scope of protection of the present patent application.
Those skilled in the art will appreciate that those matters not described in detail in the present specification are well known in the art.
Claims (6)
1. An ultra-miniature turbojet power unit, characterized in that it comprises: a single-stage radial-flow compressor, a microscale combustion chamber and a single-stage radial-inflow turbine, wherein,
the compressor rotor of the single-stage radial-flow compressor and the turbine rotor of the single-stage radial-flow turbine form an integral rotor;
the single-stage radial-flow compressor and the single-stage radial-inflow turbine form the micro-scale combustion chamber therebetween, and the micro-scale combustion chamber comprises: the high-temperature and high-pressure gas flows into a driving turbine rotor of the turbine along the single-stage radial direction to do work, and then is converted into axial discharge again;
the turbojet power device is formed by adopting 6 monocrystalline silicon wafers with the thickness of 0.4mm, 2 monocrystalline silicon wafers with the thickness of 0.6mm and 3 monocrystalline silicon wafers with the thickness of 0.8mm to be etched and formed separately and then sealing the adjacent silicon wafers into an air flow channel through a diffusion bonding connection technology.
2. The apparatus of claim 1 wherein the single stage radial compressor is comprised of an air intake, the compressor rotor, a compressor stator, and a fuel injection port; wherein the content of the first and second substances,
the air inlet channel is arranged at the center of the compressor stator;
the compressor rotor is provided with a plurality of blades, so that the airflow passing through the air inlet channel flows in the radial direction after passing through the compressor rotor and then passes through diffuser blades on the compressor stator to play a role in speed reduction and diffusion;
the fuel injection port is arranged on the section of the stator outlet of the compressor in an annular structure, and is in a uniform multipoint form in the circumferential direction.
3. The apparatus of claim 1, wherein the micro-scale combustion chamber is comprised of a ring-cavity thermal insulation sandwich, a flame tube, an ignition electric nozzle, a protruding flare, and a gas outlet, wherein,
the annular cavity heat insulation interlayer is arranged on the outer side of the micro-scale combustion chamber in an annular cylindrical shape so as to cool the chamber wall of the combustion chamber and preheat premixed gas;
the protruding flaring is arranged at the position of the airflow inlet of the flame tube, and the ignition electric nozzles are circumferentially arranged in the front of the combustion chamber so as to ignite the combustion chamber to organize and burn, and then flow into the single-stage radial inflow turbine through the gas outlet.
4. The apparatus of claim 1 wherein the single stage centripetal turbine is comprised of a turbine guide, a turbine rotor, and an exhaust outlet,
the turbine guider is provided with a plurality of guide blades, and after high-temperature airflow passes through the turbine guider, the high-temperature airflow flows into the blades of the turbine rotor along the radial direction to drive the turbine rotor to do work, and then the high-temperature airflow is bent by 90 degrees and is axially discharged from the annular exhaust passage.
5. The apparatus of claim 2, wherein the compressor stator has an outlet diameter of 7mm and the fuel injection ports are circumferentially disposed at the outlet.
6. The apparatus of claim 1, wherein the annular chamber thermal interlayer has a width of 0.4mm and an axial length of 4.4 mm.
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